The invention relates to a method for fuel injection into a staged or steped gas turbine combustion chamber with separate fuel injection nozzles for each stage, whereby at least one stage is able to be switched off for specific operating conditions by interrupting the fuel supply. Furthermore, the invention relates to a fuel injection mechanism for execution of the fuel injection method according to the invention.
For the state of the art, reference may be had to WO 95/17632 as an example.
Gas turbine combustion chambers, in particular annular combustion chambers of gas turbines, which operate with staged combustion/staged fuel injection, are increasingly gaining importance for the purpose of reducing the oxides of nitrogen. Typically, a pilot combustion chamber as well as a main combustion chamber is provided which each form constituting a so-called stage or step. Of course, further gradations/stages may also be provided in addition to these two stages. The pilot combustion chamber has as a first stage one or more pilot burners which, in the preferred case of application, comprises an annular combustion chamber and includes fuel injection nozzles in an annular arrangement; likewise, the second stage, namely the main combustion chamber, has several main burners also in the form of several injection nozzles preferably also in an annular arrangement, but optimized for reducing the oxides of nitrogen.
The attached FIG. 2 shows a basic illustration for such a staged gas turbine combustion chamber. In this case, the combustion chamber outer wall is marked with reference number 20 and the combustion chamber inner wall with reference number 21. In addition, these two walls 20, 21 are surrounded by enveloping walls 20a, 21a which also define on the left side the combustion chamber entrance 22a and on the right side the combustion chamber exit 22b. Typically, several sets of pilot and main combustion chambers such as are shown in FIG. 2 are arranged symmetrically about the center line or axis 23 in a gas turbine engine.
A separating wall structure 24 is provided within the left half of each combustion chamber. The so-called pilot combustion chamber 25a is situated between this separating wall structure 24 and the center line 23, while the so-called main combustion chamber 25b is below this separating wall structure 24, that is, radially outwardly of the pilot chamber 25a. Assigned to the pilot combustion chamber 25a are pilot nozzles 26a, while main nozzles 26b are provided for the main combustion chamber 25b. Fuel and/or an air-fuel mixture is introduced via these nozzles 26a, 26b into the combustion chambers, while a main air current 27 makes its way via the combustion chamber entrance 22a into the individual combustion chambers 25a, 25b. Furthermore, admixed air 28 can enter via openings in the outer wall 20, in the inner wall 21, as well as in the separating wall structure 24 into the individual combustion chambers 25a, 25b. The air-fuel mixture burned in the pilot burner combustion chamber 25a and/or in the main combustion chamber 25b as well as in the junction of these two combustion chambers is finally carried off via the combustion chamber exit 22b. 
Only the pilot nozzles 26a are operated in lower stress points (low load operations) of the gas turbine, that is to say, the injection nozzles of the main burner 26b are not supplied with fuel. In higher load points of the gas turbine, the main burners 26b are operated in addition to the pilot burners 26a, in such a way that their injection nozzles are then supplied with fuel. Typically, the pilot combustion chamber 25a, which is also operated singly for starting the gas turbine and for raising the engine speed up to idle, is operated throughout the entire operating performance range of the gas turbine, particularly in an airplane gas turbine, in order to create an ignition source for the main burners 26b which are only switched on when necessary. The purpose of the staged combustion lies in the minimizing of harmful substance emissions, in particular NOx. This is achieved in that the respective burner sizes can be better adapted to the given power requirement. Thus, to reduce NOx the combustion chamber temperature should be as low as possible, which can be achieved by targeted air supplying (admixed air 28) into the combustion chamber zone. In this connection, the respective stages, namely the pilot burner 26a/the main burners 26b are designed for special air-fuel ratios. In the case of low load points of the gas turbine, in which altogether only relatively little fuel is burned, the air-fuel ratio reaching the main burners 26b would be too great to be able to support a reasonable combustion at all. For this reason, the main burners 26b are switched on only in higher load conditions of the gas turbine.
FIG. 3 shows graphically the strategy according to which the individual burners, namely, the pilot burners 26a as well as the main burners 26b, are supplied with fuel in this connection. The total fuel flow for the two burners is plotted on the abscissa of this diagram, and the percentage of the pilot burners 26a and/or of the main burners 26b in this total fuel flow is plotted on the ordinate. The corresponding characteristic curve of the pilot burner 26a is marked with the letter A and that of the main burners 26b with the letter B. One recognizes that with only a slight total fuel flow at first, that is, in the left section of this diagram, only the pilot burners 26a are operated, in such a way that their share of the total fuel flow is 100%. As total fuel flow increases, the main burners 26b are then switched on, namely at the switch-on point Z. In so doing, however, there should not be a sudden power increase. Rather, a smooth power increase is desired, in such a way that with a relatively slight supply to the main burners 26b, the pilot burners 26a are supplied at the same time with a smaller fuel quantity. This switch-on point Z is therefore extremely critical with regard to its setting because there must always be a suitable air-fuel ratio in the pilot burners 26a as well as in the main burners 26b. In this regard, the same considerations also apply with respect to a reduction in or withdrawal of power of the gas turbine, that is, if the main burners 26b after being operated at first are switched off again. To avoid instabilities in the immediate surroundings of this switch-on point Z, a control that contains a hysteresis is proposed for this in WO 95/17632 mentioned above. As thrust increases, the main burners are switched on only at a higher total fuel throughput than when they are switched off as thrust decreases.
But since it is desirable to always have a defined fuel throughput in a defined load point or thrust status of the gas turbine, i.e., regardless of whether it is a matter of a thrust increase or a thrust reduction, the invention addresses the technical problem of providing another solution for the above-described problems in connection with the operation of a second stage with a first stage.
This technical problem is solved in that at least the stage which is able to be switched off can be operated with pulsed fuel injection. Appropriate fuel injection mechanisms for execution of this fuel injection method according to the invention are described in claims 5 and 6, while the further subclaims contain advantageous designs and further improvements.
The objectives of the invention are twofold: firstly, to pulse the fuel flow hence the combustion in the main combustion chamber and, secondly, to extend the operation region of the main burner stage further into the lean operating region. Pulsing the fuel flow is desirable since it is well known that pulsed combustion results in lower emissions of oxides of nitrogen.
According to the invention, at least the stage which is able to be switched off, i.e., preferably the above-mentioned main burner 25b, can be operated with pulsed fuel injection. This means that fuel injection is then not continuous but rather discontinuous. The fuel is thus introduced, practically clocked, into the combustion chamber, whereby the pulsation frequency may range from a few Hz to several 100 Hz. This pulsed injection results in a likewise pulsed combustion at least in theory. In this connection, a favorable air-fuel ratio can be set for each injection pulse or for each so-called combustion pulse. In this way, at least at low fuel quantities, fuel is no longer injected continuously but rather intermittently from then on. Thus, when favorable air-fuel ratios are set, overall, clearly less fuel can be injected than is possible with a conventional continuous injection. In particular, due to the pulsed injection the so-called switch-on point Z, can be reduced to a lower percentage of total fuel flow thus extending the operating region of the main burner stage to a lower power level. Thus, on one hand a smooth transition when switching on the second stage is attainable and, on the other hand, a defined fuel quantity is actually introduced into the combustion chamber for each operating point/thrust value, regardless of whether it is a matter of a thrust increase or a thrust reduction. The pulsation frequency, which should preferably be variable in order to be able to set a favorable combustion in a number of operating points, can preferably be above the characteristic frequencies of possible combustion chamber chugging, in such a way that no negative effects on combustion efficiency or on thrust or on noise generation need be feared. Rather, a combustion with a favorable efficiency can always be achieved, because there is a favorable air-fuel ratio for each combustion or injection pulse. Whereas, with the presently typical, continuous fuel injection into the main combustion chamber (able to be switched off), the minimum value of the fuel throughput is determined by the instability of the combustion due to too meager an air-fuel mixture. With the invention""s pulsed fuel injection for each fuel pulse, a greater air-fuel ratio is achievable, in such a way that by targeted selection of the pulsation frequency, a stable combustion or a series of stable combustion pulses is still attainable even with clearly less total fuel supplied.
As already explained, the pulsation frequency of the discontinuous fuel injection can be varied, in order to be able to adapt the total fuel quantity injected within a certain period of time to the respective operating point of the gas turbine. But it is also desirable to be able to vary the fuel quantity able to be introduced with each injection pulse, whereby there are several possibilities for this. On the one hand, with a constant fuel quantity the injection period per unit of time can be altered, and on the other hand, with a constant injection period the fuel quantity introduced during this period can be altered. Of course, it is also possible to combine these two strategies, just as the pulsation frequency can additionally be adapted, in such a way that altogether, the optimal fuel injection in each case can be selected by means of the many variation possibilities for each operating point of the gas turbine. Thus, the fuel can be controlled by pulse width modulation. In this connection, it should be pointed out that in the high load operating conditions, one can switch from the pulsed injection to a continuous fuel injection, of course.
In addition, a further advantage of the pulsed fuel injection should be pointed out. Due to the targeted selection of the pulsation frequency, namely the typical combustion frequencies can be controlled in such a way that the so-called xe2x80x9ccombustion hummingxe2x80x9d, which can occur with unstable combustion and with low fuel throughput and results from the characteristic frequencies of possible combustion chamber chugging, can be minimized. It should also be pointed out that the first stage or pilot combustion chamber, which is usually not switched off, can or should preferably operate with a continuous fuel injection, in particular also in order to ensure reliable ignition of the air-fuel mixture in the second stage or main combustion chamber.
An advantageous fuel injection mechanism for execution of such a pulsed fuel injection can comprise an electromagnetically and/or hydraulically actuated fuel injection valve the time of opening and open period of which can be adjusted in a targeted manner. Such fuel injection valves are known from reciprocating internal combustion engines. Appropriately modified, such fuel injection valves can then be used either to directly inject the fuel into the combustion chamber of a gas turbine or they can be connected upstream from an essentially typical fuel injection nozzle.
A fuel injection mechanism for execution of a pulsed fuel injection according to the invention can consist of a suitable pulsation control valve and a dosing valve that is arranged upstream from a basically typical fuel injection nozzle ending in the combustion chamber. In addition to the pulsation control valve, a dosing valve can be arranged upstream from this injection nozzle, whereby it is particularly advantageous to combine the pulsation control valve and the dosing valve in a component hereinafter designated a xe2x80x9cpulse-doserxe2x80x9d.